Methods and apparatus for cooling turbine engine combustor exit temperatures

ABSTRACT

A method facilitates assembling a combustor for a gas turbine engine. The method comprises coupling an inner liner to an outer liner such that a combustion chamber is defined therebetween, positioning an outer support a distance radially outward from the outer liner, and positioning an inner support a distance radially inward from the inner liner. The method also comprises forming at least two rows of impingement openings extending through at least one of the inner support and the outer support for channeling impingement cooling air therethrough towards at least one of the inner liner and the outer liner, and forming at least one row of dilution openings extending through at least one of the inner liner and the outer liner for channeling dilution cooling air therethrough into the combustion chamber

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

The U.S. Government may have certain rights in this invention pursuantto contract number DAAE07-00-C-N086.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines, moreparticularly to combustors used with gas turbine engines.

Known turbine engines include a compressor for compressing air which issuitably mixed with a fuel and channeled to a combustor wherein themixture is ignited for generating hot combustion gases. At least someknown combustors include an inner liner that is coupled to an outerliner such that a combustion chamber is defined therebetween.Additionally, an outer support is coupled radially outward from theouter liner such that an outer cooling passage is defined therebetween,and an inner support is coupled radially inward from the inner linersuch that an inner cooling passage is defined therebetween.

Within at least some known recuperated gas turbine engines, coolingrequirements of turbines may create a pattern factor requirement at thecombustor that may be difficult to achieve because of combustor designcharacteristics associated with recuperated gas turbine engines. Morespecifically, because of space considerations, such combustors may beshorter than other known gas turbine engine combustors. In addition, incomparison to other known gas turbine combustors, such combustors mayinclude a steeply angled flowpath and large fuel injector spacing.

Accordingly, at least some known combustors include a dilution patternof a single row of dilution jets to facilitate controlling the combustorexit temperatures. The dilution jets are supplied cooling air from animpingement array of openings extending through the inner and outersupports. However, because of cooling considerations downstream from thecombustor and because of the limited number and relative orientation ofsuch impingement and dilution openings, such combustors may only receiveonly limited dilution air from such openings.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a method for assembling a combustor for a gas turbineengine is provided. The method comprises coupling an inner liner to anouter liner such that a combustion chamber is defined therebetween,positioning an outer support a distance radially outward from the outerliner, and positioning an inner support a distance radially inward fromthe inner liner. The method also comprises forming at least two rows ofimpingement openings extending through at least one of the inner supportand the outer support for channeling impingement cooling airtherethrough towards at least one of the inner liner and the outerliner, and forming at least one row of dilution openings extendingthrough at least one of the inner liner and the outer liner forchanneling dilution air therethrough into the combustion chamber.

In another aspect, a combustor for a gas turbine engine is provided. Thecombustor includes an inner liner, an outer liner, an outer support, andan inner support. The outer liner is coupled to the inner liner todefine a combustion chamber therebetween. The outer support is radiallyoutward from the outer liner such that an outer passageway is definedbetween the outer support and the outer liner. The inner support isradially inward from the inner liner such that an inner passageway isdefined between the inner support and the inner liner. At least one ofthe inner support and the outer support includes at least two rows ofimpingement openings arranged in an array and extending therethrough forchanneling impingement cooling air towards at least one of the innerliner and the outer liner. At least one of the inner liner and the outerliner includes at least one row of dilution openings extendingtherethrough for channeling dilution air into the combustion chamber.

In a further aspect, a gas turbine engine including a combustor isprovided. The combustor includes at least one injector, an inner liner,an outer liner, an outer support, and an inner support. The inner lineris coupled to the outer liner to define a combustion chambertherebetween. The inner and outer liners further define an injectoropening, and the injector extends substantially concentrically throughthe injector opening. The outer support is spaced radially outward fromthe outer liner. The inner support is spaced radially inward from theinner liner. At least one of the inner support and the outer supportincludes at least two rows of impingement openings arranged in an arrayand extending therethrough for channeling impingement cooling airtowards at least one of the inner liner and the outer liner. At leastone of the inner liner and the outer liner includes at least one row ofdilution openings extending therethrough for channeling dilution airinto the combustion chamber.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic of a gas turbine engine.

FIG. 2 is a cross-sectional illustration of a portion of an annularcombustor used with the gas turbine engine shown in FIG. 1;

FIG. 3 is a roll-out schematic view of a portion of the combustor shownin FIG. 2 and taken along area 3;

FIG. 4 is a roll-out schematic view of a portion of the combustor shownin FIG. 2 and taken along area 4.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a schematic illustration of a gas turbine engine 10 includinga compressor 14, and a combustor 16. Engine 10 also includes a highpressure turbine 18 and a low pressure turbine 20. Compressor 14 andturbine 18 are coupled by a first shaft 24, and turbine 20 drives asecond output shaft 26. Shaft 26 provides a rotary motive force to drivea driven machine, such as, but, not limited to a gearbox, atransmission, a generator, a fan, or a pump. Engine 10 also includes arecuperator 28 that has a first fluid path 29 coupled serially betweencompressor 14 and combustor 16, and a second fluid path 31 that isserially coupled between turbine 20 and ambient 35. In one embodiment,the gas turbine engine is an LV100 engine available from GeneralElectric Company, Cincinnati, Ohio. In the exemplary embodiment,compressor 14 is coupled by a first shaft 24 to turbine 18, andpowertrain and turbine 20 are coupled by a second shaft 26.

In operation, air flows through high pressure compressor 14. The highlycompressed air is delivered to recouperator 28 where hot exhaust gasesfrom turbine 20 transfer heat to the compressed air. The heatedcompressed air is delivered to combustor 16. Airflow from combustor 16drives turbines 18 and 20 and passes through recouperator 28 beforeexiting gas turbine engine 10. In the exemplary embodiment, duringoperation, air flows through compressor 14, and the highly compressedrecuperated air is delivered to combustor 16.

FIG. 2 is a cross-sectional illustration of a portion of an annularcombustor 16. FIG. 3 is a roll-out schematic view of a portion ofcombustor 16 and taken along area 3 (shown in FIG. 2). FIG. 4 is aroll-out schematic view of a portion of combustor 16 and taken alongarea 4 (shown in FIG. 2). Combustor 16 includes an annular outer liner40, an outer support 42, an annular inner liner 44, an inner support 46,and a dome 48 that extends between outer and inner liners 40 and 44,respectively.

Outer liner 40 and inner liner 44 extend downstream from dome 48 anddefine a combustion chamber 54 therebetween. Combustion chamber 54 isannular and is spaced radially inward between liners 40 and 44. Outersupport 42 is coupled to outer liner 40 and extends downstream from dome48. Moreover, outer support 42 is spaced radially outward from outerliner 40 such that an outer cooling passageway 58 is definedtherebetween. Inner support 46 also is coupled to, and extendsdownstream from, dome 48. Inner support 46 is spaced radially inwardfrom inner liner 44 such that an inner cooling passageway 60 is definedtherebetween.

Outer support 42 and inner support 46 are spaced radially within acombustor casing 62. Combustor casing 62 is generally annular andextends around combustor 16. More specifically, outer support 42 andcombustor casing 62 define an outer passageway 66 and inner support 42and combustor casing 62 define an inner passageway 68. Outer and innerliners 40 and 44 extend to a turbine nozzle 69 that is downstream fromliners 40 and 44.

Combustor 16 also includes a dome assembly 70 which includes an airswirler 90. Specifically, air swirler 90 extends radially outwardly andupstream from a dome plate 72 to facilitate atomizing and distributingfuel from a fuel nozzle 82. When fuel nozzle 82 is coupled to combustor16, nozzle 82 circumferentially contacts air swirler 90 to facilitateminimizing leakage to combustion chamber 54 between nozzle 82 and airswirler 90.

Combustor dome plate 72 is mounted upstream from outer and inner liners40 and 44, respectively. Dome plate 72 contains a plurality ofcircumferentially spaced air swirlers 90 that extend through dome plate72 into combustion chamber 54 and each include a center longitudinalaxis of symmetry 76 that extends therethrough. Fuel is supplied tocombustor 16 through a fuel injection assembly 80 that includes aplurality of circumferentially-spaced fuel nozzles 82 that extendthrough air swirlers 90 into combustion chamber 54. More specifically,fuel injection assembly 80 is coupled to combustor 16 such that eachfuel nozzle 82 is substantially concentrically aligned with respect toair swirlers 90, and such that nozzle 82 extends downstream into airswirler 90. Accordingly, a centerline 84 extending through each fuelnozzle 82 is substantially co-linear with respect to air swirler axis ofsymmetry 76.

Because of the steeply angled flowpath 100 defined within combustor 16,circumferential spacing between adjacent fuel nozzles 82 and airswirlers 90, and downstream component cooling requirements, combustiongases generated within combustor 16 are cooled prior to being dischargedfrom combustor 16 to enable combustor 16 to maintain a pre-determinedpattern factor. Combustor pattern factor is generally defined as:PF=(T4_(peak) −T4_(avg))/(T4_(avg) −T35)where T4 refers to the combustor exit temperature, T35 refers to thecombustor inlet temperature, and T4 peak refers to the maximumtemperature measured, and T4_(avg.) refers to the average of thetemperatures measured. Pattern factor is a measure of the distortion incombustor exit temperature and generally, a lower value is moredesirable.

Accordingly, combustor outer and inner liners 40 and 44, each include aplurality of dilution jets 110 to facilitate locally cooling combustiongases generated within combustion chamber 54, and to provide radial andcircumferential exit temperature distribution. In the exemplaryembodiment, dilution jets 110 are substantially circular and extendthrough liners 40 and 44. More specifically, outer liner 40 includes aplurality of primary larger diameter dilution openings 120, a pluralityof smaller diameter dilution openings 122, and a plurality of secondarydilution openings 124. Openings 120, 122, and 124 extendcircumferentially around combustor 16.

Smaller diameter outer primary dilution openings 122 are positionedsubstantially axially downstream with respect to air swirler centerline76 at pre-determined distances D₁ downstream from dome 72. Morespecifically, in the exemplary embodiment, smaller outer primarydilution openings 122 are positioned downstream from dome plate 72 at adistance D₁ that is approximately equal 0.65 combustor passage heightsh₁. Combustor passage heights h₁ is defined as the measured distancebetween outer and inner liners 40 and 44 at combustor chamber upstreamend 74.

Larger diameter outer primary dilution openings 120 have a largerdiameter d₂ than a diameter d₃ of smaller diameter outer primarydilution openings 122, and are positioned between adjacent air swirlers90 at the same axial locations as openings 122. In one embodiment,larger diameter openings 120 have a diameter d₂ that is approximatelyequal 0.307 inches, and smaller diameter openings 122 have a diameter d₃that is approximately equal 0.243 inches. Accordingly, each opening 120is between a pair of circumferentially adjacent openings 122.

Outer secondary dilution openings 124 each have a diameter d₄ that issmaller than that of openings 120 and 122, and are each located at apredetermined axial distance D₅ aft of openings 120 and 122. In oneembodiment, openings 124 have a diameter d₄ that is approximately equal0.168 inches. More specifically, in the exemplary embodiment, openings124 are approximately 0.25 passage heights h₁ downstream from openings120 and 122. In addition, each secondary dilution opening 124 ispositioned downstream from, and between, a pair of circumferentiallyadjacent primary dilution openings 120 and 122.

Inner liner 44 also includes a plurality of dilution jets 110 extendingtherethrough. More specifically, inner liner 44 includes a plurality ofinner primary dilution openings 130 which each have a diameter d₆ thatis smaller than a diameter d₂ and d₃ of respective outer primarydilution openings 120 and 122. In one embodiment, openings 130 have adiameter d₆ that is approximately equal 0.228 inches. Each inner primarydilution opening 130 is circumferentially aligned with each outersecondary dilution opening 124 and between adjacent outer primarydilution openings 120 and 122. More specifically, in the exemplaryembodiment, inner primary dilution openings 130 are positioneddownstream from dome plate 72 at a distance D₈ that is approximatelyequal 0.70 combustor passage heights h₁. Accordingly, because primarydilution jets 120 and 122, and 130 are not opposed, enhanced mixing andenhanced circumferential coverage is obtained between dilution jets 110and mainstream combustor flow. Accordingly, the enhanced mixingfacilitates reducing combustor exit temperature distortion and, thusreduces pattern factor.

A number of dilution jets 110 is variably selected to facilitateachieving a desired radial and circumferential exit temperaturedistribution from combustor 16. More specifically, combustor 16 includesan equal number of outer primary dilution openings 120 and 122, outersecondary dilution openings 124, and inner primary dilution openings130. In the exemplary embodiment, combustor 16 includes eighteen largerdiameter outer primary dilution openings 120, eighteen smaller diameterouter primary dilution openings 122, and thirty-six inner primarydilution openings 130. More specifically, the number of outer primarydilution openings 120 and 122, outer secondary dilution openings 124 isselected to be twice the number of fuel injectors 82 fueling combustor16.

Outer primary dilution openings 120 and 122, and outer secondarydilution openings 124 receive air discharged through impingementopenings or jets 140 formed within outer support 42. Specifically,openings 140 are arranged in an array 144 that facilitates maximizingthe cooling airflow available for impingement cooling of outer liner 40.Within array 144, openings 140 extend circumferentially around outersupport 42, but do not extend into pre-designated interruption areas 146defined across outer support 42. More specifically, each interruptionarea 146 is formed radially outward from outer primary dilution openings120 and 122, and outer secondary dilution openings 124 to facilitateavoiding variable interaction between impingement and dilution jets 140and 110, respectively, either by entrainment or by ejector effect.

Similarly, inner primary dilution openings 130 receive air dischargedthrough impingement jets or openings 140 formed within inner support 46.Specifically, opening array 144 facilitates maximizing the coolingairflow available for impingement cooling of inner liner 44. Withinarray 144, openings 140 extend circumferentially across inner support46, but do not extend into pre-designated interruption areas 150 definedacross support 46. More specifically, each interruption area 150 isformed radially outward from inner primary dilution openings 130 tofacilitate avoiding variable interaction between impingement anddilution jets 140 and 110, respectively, either by entrainment or byejector effect.

Impingement jets 140 also supply airflow to multi-hole film coolingopenings 160 formed within outer and inner liners 40 and 44,respectively. More specifically, openings 160 are oriented to dischargecooling air therethrough for film cooling liners 40 and 44. Accordingly,the number of impingement jets 140 is selected to facilitate maximizingthe amount of cooling airflow supplied to liners 40 and 44. In theexemplary embodiment, the number of impingement jets 140 is a multipleof the number of dilution jets 110. More specifically, the number ofimpingement jets 140 and dilution jets 110 are selected to ensure thatthe pressure differential across impingement holes 140 in outer andinner supports 42 and 46, respectively, approximately matches thepressure differential across the film cooling openings 160 and acrossdilution openings 120, 122, 124, and 130.

During operation, impingement cooling air is directed throughimpingement jets 140 towards outer and inner liners 40 and 44,respectively, for impingement cooling of liners 40 and 44. The coolingair is also channeled through dilution jets 110 and through film coolingopenings 160 into combustion chamber 54. More specifically, airflowdischarged from openings 160 facilitates film cooling of liners 40 and44 such that an operating temperature of each is reduced. Airflowentering chamber 54 through jets 110 facilitates radially andcircumferentially cooling a temperature of the combustor flow path suchthat a desired exit temperature distribution is obtained. As such, thereduced combustor operating temperatures facilitate extending a usefullife of combustor 16 and the desired exit temperature distributionfacilitates extending a useful life to turbine hardware downstream ofcombustor 16.

The above-described dilution and impingement jets provide acost-effective and reliable means for operating a combustor. Morespecifically, each support includes a plurality of impingement jets thatchannel cooling air radially inward for impingement cooling of thecombustor outer and inner liners. The outer and inner liners eachinclude a plurality of dilution jets and film cooling openings whichchannel air inward into the combustion chamber. As a result, at leastsome of the impingement cooling air film cools the liners, and theremaining impingement cooling air is directed inward to facilitateradially and circumferentially cooling the combustor flow path such thata desired exit temperature distribution is obtained.

An exemplary embodiment of a combustion system is described above indetail. The combustion system components illustrated are not limited tothe specific embodiments described herein, but rather, components ofeach combustion system may be utilized independently and separately fromother components described herein. For example, the impingement jetsand/or dilution jets may also be used in combination with other enginecombustion systems.

While the invention has been described in terms of various specificembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theclaims.

1. A method for assembling a combustor for a gas turbine engine, saidmethod comprising: coupling an inner liner to an outer liner such that acombustion chamber is defined therebetween; positioning an outer supporta distance radially outward from the outer liner; positioning an innersupport a distance radially inward from the inner liner; forming atleast two rows of impingement openings extending through at least one ofthe inner support and the outer support for channeling impingementcooling air therethrough towards at least one of the inner liner and theouter liner; and forming at least one row of dilution openings extendingthrough at least one of the inner liner and the outer liner forchanneling dilution cooling air therethrough into the combustionchamber.
 2. A method in accordance with claim 1 wherein forming at leastone row of dilution openings further comprises: forming a row of firstprimary dilution openings that each have a first diameter; and forming arow of second primary dilution openings that each have a second diameterthat is larger than the first diameter of the first primary dilutionopenings.
 3. A method in accordance with claim 2 wherein forming a rowof second primary dilution openings further comprises forming the row ofsecond primary dilution openings such that each of the second primarydilution openings is between a pair of adjacent first primary dilutionopenings.
 4. A method in accordance with claim 1 further comprisingforming a plurality of film cooling openings extending through at leastone of said inner liner and said outer liner for channeling cooling airfor film cooling of at least one of said inner liner and said outerliner, wherein the plurality of film cooling openings are in flowcommunication with the at least two rows of impingement openings.
 5. Amethod in accordance with claim 4 wherein forming at least one row ofdilution openings further comprises forming the dilution openings suchthat a pressure differential across the at least two rows impingementopenings is substantially equal to a pressure differential across the atleast one row of dilution openings and said plurality of film coolingopenings.
 6. A combustor for a gas turbine engine, said combustorcomprising: an inner liner; an outer liner coupled to said inner linerto define a combustion chamber therebetween; an outer support radiallyoutward from said outer liner such that an outer passageway is definedbetween said outer support and said outer liner; and an inner supportradially inward from said inner liner such that an inner passageway isdefined between said inner support and said inner liner, at least one ofsaid inner support and said outer support comprising at least two rowsof impingement openings arranged in an array and extending therethroughfor channeling impingement cooling air towards at least one of saidinner liner and said outer liner, at least one of said inner liner andsaid outer liner comprising at least one row of dilution openingsextending therethrough for channeling dilution cooling air into saidcombustion chamber.
 7. A combustor in accordance with claim 6 whereinsaid at least one row of dilution openings facilitate radially andcircumferentially reducing exit flow temperatures from said combustor.8. A combustor in accordance with claim 6 wherein said at least one rowof dilution openings further comprises a row of first primary dilutionopenings having a first diameter, and a row of second primary dilutionopenings having a second diameter that is larger than said firstdiameter.
 9. A combustor in accordance with claim 8 wherein saidcombustor comprises an equal number of said first primary dilutionopenings and said second primary dilution openings.
 10. A combustor inaccordance with claim 8 wherein each said second primary dilutionopening is between a pair of adjacent said first primary dilutionopenings.
 11. A combustor in accordance with claim 8 wherein at leastone of said inner liner and said outer liner further comprises aplurality of film cooling openings extending therethrough for channelingcooling air for film cooling of at least one of said inner liner andsaid outer liner.
 12. A combustor in accordance with claim 11 wherein apressure differential across said at least two rows impingement openingsis substantially equal to a pressure differential across said at leastone row of dilution openings and said plurality of film coolingopenings.
 13. A gas turbine engine comprising a combustor comprising atleast one injector, an inner liner, an outer liner, an outer support,and an inner support, said inner liner coupled to said outer liner todefine a combustion chamber therebetween, said inner and outer linersfurther defining a dome opening, said injector extending substantiallyconcentrically through said dome opening, said outer support spacedradially outward from said outer liner, said inner support spacedradially inward from said inner liner, at least one of said innersupport and said outer support comprising at least two rows ofimpingement openings arranged in an array and extending therethrough forchanneling impingement cooling air towards at least one of said innerliner and said outer liner, at least one of said inner liner and saidouter liner comprising at least one row of dilution openings extendingtherethrough for channeling dilution cooling air into said combustionchamber.
 14. A gas turbine engine in accordance with claim 13 whereinsaid combustor at least one row of dilution openings facilitate radiallyand circumferentially controlling distortion in exit flow temperaturesfrom said combustor.
 15. A gas turbine engine in accordance with claim14 wherein a number of said combustor first primary dilution openings isequal to a number of said combustor second primary dilution openings.16. A gas turbine engine in accordance with claim 14 wherein saidcombustor at least one row of dilution openings further comprises a rowof first primary dilution and a row of second primary dilution openings,each of said first primary dilution openings has a first diameter thatis smaller than a second diameter of each of said second primarydilution openings.
 17. A gas turbine engine in accordance with claim 16wherein each said combustor second primary dilution opening is between apair of adjacent said first primary dilution openings.
 18. A gas turbineengine in accordance with claim 16 wherein said combustor furthercomprises a plurality of air swirlers, each said combustor first primarydilution opening is aligned downstream from, and substantially axiallyalong a centerline of, each said air swirler.
 19. A gas turbine enginein accordance with claim 14 wherein at least one of said inner liner andsaid outer liner further comprises a plurality of film cooling openingsfor channeling cooling air therethrough for film cooling at least one ofsaid inner liner and said outer liner.
 20. A gas turbine engine inaccordance with claim 19 wherein a pressure differential across saidcombustor array of impingement openings is substantially equal to apressure differential across said at least one row of dilution openingsand said plurality of film cooling openings.